U.S. patent application number 13/362330 was filed with the patent office on 2013-08-01 for buffer system that communicates buffer supply air to one or more portions of a gas turbine engine.
The applicant listed for this patent is Peter M. Munsell, Philip S. Stripinis. Invention is credited to Peter M. Munsell, Philip S. Stripinis.
Application Number | 20130192238 13/362330 |
Document ID | / |
Family ID | 48869052 |
Filed Date | 2013-08-01 |
United States Patent
Application |
20130192238 |
Kind Code |
A1 |
Munsell; Peter M. ; et
al. |
August 1, 2013 |
BUFFER SYSTEM THAT COMMUNICATES BUFFER SUPPLY AIR TO ONE OR MORE
PORTIONS OF A GAS TURBINE ENGINE
Abstract
A gas turbine engine includes a buffer system that communicates
a buffer supply air to a portion of the gas turbine engine. The
buffer system includes a first bleed air supply having a first
pressure, a second bleed air supply having a second pressure that
is greater than the first pressure, and a valve that selects
between the first bleed air supply and the second bleed air supply
to communicate the buffer supply air to the portion of the gas
turbine engine.
Inventors: |
Munsell; Peter M.; (Granby,
CT) ; Stripinis; Philip S.; (Rocky Hill, CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Munsell; Peter M.
Stripinis; Philip S. |
Granby
Rocky Hill |
CT
CT |
US
US |
|
|
Family ID: |
48869052 |
Appl. No.: |
13/362330 |
Filed: |
January 31, 2012 |
Current U.S.
Class: |
60/772 ;
60/785 |
Current CPC
Class: |
Y02T 50/676 20130101;
Y02T 50/60 20130101; F05D 2260/20 20130101; F02C 7/18 20130101;
F01D 25/183 20130101; F01D 25/125 20130101; F02C 6/08 20130101;
F05D 2270/335 20130101 |
Class at
Publication: |
60/772 ;
60/785 |
International
Class: |
F02C 6/08 20060101
F02C006/08; F02C 7/12 20060101 F02C007/12 |
Claims
1. A gas turbine engine, comprising: a buffer system that
communicates a buffer supply air to a portion of the gas turbine,
wherein said buffer system includes: a first bleed air supply
having a first pressure; a second bleed air supply having a second
pressure that is greater than said first pressure; and a valve that
selects between said first bleed air supply and said second bleed
air supply to communicate said buffer supply air to said portion of
the gas turbine engine.
2. The gas turbine engine as recited in claim 1, wherein said
portion is at least one bearing compartment of the gas turbine
engine.
3. The gas turbine engine as recited in claim 1, comprising a
controller that selectively modulates said valve in response to a
power condition.
4. The gas turbine engine as recited in claim 3, comprising a
sensor that detects said power condition.
5. The gas turbine engine as recited in claim 1, wherein said first
bleed air supply is sourced from a location of the gas turbine
engine that is upstream from a source of said second bleed air
supply.
6. The gas turbine engine as recited in claim 1, wherein first
bleed air supply is communicated as said buffer supply air in
response to a high power condition of the gas turbine engine and
said second bleed air supply is communicated as said buffer supply
air in response to a low power condition of the gas turbine
engine.
7. A gas turbine engine, comprising: a compressor section; a
combustor in fluid communication with said compressor section; a
turbine section in fluid communication with said combustor; at
least one shaft that interconnects at least a portion of said
compressor section and said turbine section; a bearing structure
that supports said at least one shaft, wherein said bearing
structure includes a bearing compartment; and a buffer system that
selectively communicates a buffer supply air to pressurize said
bearing compartment, wherein a low pressure bleed air supply is
communicated to said bearing compartment in response to a high
power condition of the gas turbine engine and a high pressure bleed
supply air is communicated to said bearing compartment in response
to a low power condition of the gas turbine engine.
8. The gas turbine engine as recited in claim 7, wherein said
buffer system includes a valve that selectively modulates between
said low pressure bleed air supply and said high pressure bleed air
supply.
9. The gas turbine engine as recited in claim 7, wherein said high
power condition includes one of a take-off condition, a climb
condition and a cruise condition.
10. The gas turbine engine as recited in claim 7, wherein said low
power condition includes one of a ground condition, an idle
condition and descent idle.
11. The gas turbine engine as recited in claim 7, comprising a fan
section driven by a geared architecture.
12. The gas turbine engine as recited in claim 7, wherein the gas
turbine engine is a high bypass geared aircraft engine having a
bypass ratio of greater than about six (6).
13. The gas turbine engine as recited in claim 11, wherein the gas
turbine engine includes a low Fan Pressure Ratio of less than about
1.45.
14. The gas turbine engine as recited in claim 7, wherein the
buffer system includes a first circuit that provides the low
pressure bleed air supply to the bearing compartment and a second
circuit that provides the high pressure bleed supply air to the
bearing compartment.
15. The gas turbine engine as recited in claim 7, wherein said
bearing compartment comprises a first bearing compartment and a
second bearing compartment.
16. A method of cooling a portion of a gas turbine engine,
comprising: selecting between a first bleed air supply and a second
bleed air supply that are supplied to a valve; communicating the
first bleed air supply to cool the portion in response to a low
power condition of the gas turbine engine; and communicating the
second bleed air supply to cool the portion in response to a high
power condition of the gas turbine engine.
17. The method as recited in claim 16, wherein the first bleed air
supply is a high pressure bleed air supply and the second bleed air
supply is a low pressure bleed air supply.
18. The method as recited in claim 16, comprising the step of:
identifying a power condition of the gas turbine engine prior to
the steps of communicating.
19. The method as recited in claim 16, wherein the low power
condition includes one of a ground condition, an idle condition and
a descent idle condition.
20. The method as recited in claim 16, wherein the high power
condition includes one of a take-off condition, a climb condition
and a cruise condition.
21. The gas turbine engine as recited in claim 1, wherein said
valve is in fluid communication with each of said first bleed air
supply and said second bleed air supply.
22. The gas turbine engine as recited in claim 1, wherein said
first bleed air supply and said second bleed air supply are each
sourced from a high pressure compressor of the gas turbine
engine.
23. The gas turbine engine as recited in claim 22, wherein said
first bleed air supply is sourced from an upstream stage of said
high pressure compressor and said second bleed air supply is
sourced from one of a middle stage and a downstream stage of said
high pressure compressor.
24. The gas turbine engine as recited in claim 1, wherein said
buffer supply air is communicated from said valve to at least one
bearing compartment located upstream from a source of said first
bleed air supply and said second bleed air supply.
25. The gas turbine engine as recited in claim 1, wherein said
valve is a passive valve that uses only a single input that is
directly measured to switch between said first bleed air supply and
said second bleed air supply.
26. The gas turbine engine as recited in claim 8, wherein said
valve is in fluid communication with each of said low pressure
bleed air supply and said high pressure bleed air supply.
27. The method as recited in claim 16, wherein the step of
selecting between the first bleed air supply and the second bleed
air supply is performed prior to the steps of communicating.
28. The method as recited in claim 16, wherein the portion includes
multiple bearing compartments.
29. The method as recited in claim 16, wherein at least one of the
steps of communicating includes communicating either the first
bleed air supply or the second bleed air supply to a location that
is upstream from a source of the first bleed air supply and the
second bleed air supply.
30. The method as recited in claim 16, comprising the steps of:
prior to the steps of communicating, sourcing the first bleed air
supply from one of a middle stage and a downstream stage of a high
pressure compressor and sourcing the second bleed air supply from
an upstream stage of the high pressure compressor.
Description
BACKGROUND
[0001] This disclosure relates to a gas turbine engine, and more
particularly to a buffer system that can communicate a buffer
supply air to one or more portions of the gas turbine engine.
[0002] Gas turbine engines typically include at least a compressor
section, a combustor section and a turbine section. During
operation, air is pressurized in the compressor section and is
mixed with fuel and burned in the combustor section to generate hot
combustion gases. The hot combustion gases are communicated through
the turbine section which extracts energy from the hot combustion
gases to power the compressor section and other gas turbine engine
modes.
[0003] Gas turbine engines typically include shafts that support a
plurality of airfoil supporting rotors of the compressor section
and the turbine section. Generally, these shafts are supported by
bearing structures that define bearing compartments. The bearing
compartments house one or more bearings and contain lubricant that
is used to lubricate the bearings. The lubricant is contained
within the bearing compartment by one or more seals. A
predetermined differential pressure must be maintained across the
seals so the lubricant cannot leak past the seals.
SUMMARY
[0004] A gas turbine engine includes a buffer system that
communicates a buffer supply air to a portion of the gas turbine
engine. The buffer system includes a first bleed air supply having
a first pressure, a second bleed air supply having a second
pressure that is greater than the first pressure, and a valve that
selects between the first bleed air supply and the second bleed air
supply to communicate the buffer supply air to the portion of the
gas turbine engine.
[0005] In a further embodiment of the foregoing gas turbine engine
embodiment, the portion can include at least one bearing
compartment of the gas turbine engine.
[0006] In a further embodiment of either of the foregoing gas
turbine engine embodiments, a controller can selectively modulate
the valve in response to a power condition.
[0007] In a further embodiment of any of the foregoing gas turbine
engine embodiments, a sensor can detect the power condition.
[0008] In a further embodiment of any of the foregoing gas turbine
engine embodiments, the first bleed air supply can be sourced from
a location of the gas turbine engine that is upstream from a source
of the second bleed air supply.
[0009] In a further embodiment of any of the foregoing gas turbine
engine embodiments, the first bleed air supply can be communicated
as the buffer supply air in response to a high power condition of
the gas turbine engine and the second bleed air supply can be
communicated as the buffer supply air in response to a low power
condition of the gas turbine engine.
[0010] In another exemplary embodiment, a gas turbine engine
includes a compressor section, a combustor in fluid communication
with the compressor section, a turbine section in fluid
communication with the combustor, at least one shaft that
interconnects the portion of the compressor section and the turbine
section, and a bearing structure that supports the at least one
shaft. The bearing structure can include a bearing compartment. A
buffer system can selectively communicate a buffer supply air to
pressurize the bearing compartment. A low pressure bleed air supply
is communicated to the bearing compartment in response to the high
power condition of the gas turbine engine and a high pressure bleed
supply air is communicated to the bearing compartment in response
to a low power condition of the gas turbine engine.
[0011] In a further embodiment of the foregoing gas turbine engine
embodiment, the buffer system can include a valve that selectively
modules between the low pressure bleed air supply and the high
pressure bleed air supply.
[0012] In a further embodiment of either of the foregoing gas
turbine engine embodiments, the high power condition can include
one of a take-off condition, a climb condition and a cruise
condition.
[0013] In a further embodiment of any of the foregoing gas turbine
engine embodiments, the low power condition can include one of a
ground condition, an idle condition and descent idle.
[0014] In a further embodiment of any of the foregoing gas turbine
engine embodiments, a fan section can be driven by a geared
architecture.
[0015] In a further embodiment of any of the foregoing gas turbine
engine embodiments, the gas turbine engine can include a high
bypass geared aircraft engine having a bypass ratio of greater than
about six (6).
[0016] In a further embodiment of any of the foregoing gas turbine
engine embodiments, the gas turbine engine includes a low fan
pressure ratio of less than about 1.45.
[0017] In a further embodiment of any of the foregoing gas turbine
engine embodiments the buffer system includes a first circuit that
provides the low pressure bleed air supply to the bearing
compartment and a second circuit that provides the high pressure
bleed supply air to the bearing compartment.
[0018] In a further embodiment of any of the foregoing gas turbine
engine embodiments the bearing compartment comprises a first
bearing compartment and a second bearing compartment.
[0019] In yet another exemplary embodiment, a method of cooling a
portion of a gas turbine engine includes communicating a first
bleed air supply to the portion in response to a low power
condition of the gas turbine engine. A second bleed air supply can
be communicated to the portion in response to a high power
condition of the gas turbine engine.
[0020] In a further embodiment of the foregoing method embodiment,
the first bleed air supply can be a high pressure bleed air supply
and the second bleed air supply can be a low pressure bleed air
supply.
[0021] In a further embodiment of either of the foregoing method
embodiments, a power condition of the gas turbine engine is
identified prior to the steps of communicating.
[0022] In a further embodiment of any of the foregoing method
embodiments, the low power condition can include one of a ground
condition, an idle condition and descent idle.
[0023] In a further embodiment of any of the foregoing method
embodiments, the high power condition includes one of a take-off
condition, a climb condition and a cruise condition.
[0024] The various features and advantages of this disclosure will
become apparent to those skilled in the art from the following
detailed description. The drawings that accompany the detailed
description can be briefly described as follows.
BRIEF DESCRIPTION OF THE DRAWINGS
[0025] FIG. 1 schematically illustrates a cross-sectional view of a
gas turbine engine.
[0026] FIG. 2 illustrates a schematic cross-section of a portion of
the gas turbine engine.
[0027] FIG. 3 illustrates an example buffer system that can be
incorporated into a gas turbine engine.
[0028] FIG. 4 illustrates another example buffer system that can be
incorporated into a gas turbine engine.
[0029] FIG. 5 illustrates yet another example buffer system that
can be incorporated into a gas turbine engine.
DETAILED DESCRIPTION
[0030] FIG. 1 schematically illustrates a gas turbine engine 20.
The gas turbine engine 20 disclosed herein is a two spool turbofan
engine that generally incorporates a fan section 22, a compressor
section 24, a combustor section 26 and a turbine section 28.
Alternative engines might include an augmenter section (not shown)
among other systems or features. The fan section 22 drives air
along a bypass flow path B while the compressor section 24 drives
air along a core flow path C for compression and communication into
the combustor section 26. The hot combustion gases generated in the
combustor section 26 are expanded through the turbine section 28.
Although depicted as a turbofan gas turbine engine in the disclosed
non-limiting embodiment, it should be understood that the concepts
described herein are not limited to turbofan engines and these
teachings could extend to other types of turbine engines, including
but not limited to three spool engine architectures.
[0031] The gas turbine engine 20 generally includes a low speed
spool 30 and a high speed spool 32 mounted for rotation about an
engine centerline longitudinal axis A relative to an engine static
structure 36 via several bearing structures 38. It should be
understood that various bearing structures 38 at various locations
may alternatively or additionally be provided.
[0032] The low speed spool 30 generally includes an inner shaft 40
that interconnects a fan 42, a low pressure compressor 44 and a low
pressure turbine 46. The inner shaft 40 can be connected to the fan
42 through a geared architecture 48 to drive the fan 42 at a lower
speed than the low speed spool 30. The high speed spool 32 includes
an outer shaft 50 that interconnects a high pressure compressor 52
and a high pressure turbine 54. In this example, the inner shaft 40
and the outer shaft 50 are supported at a plurality of points by
bearing structures 38 positioned within the engine static structure
36. In one non-limiting embodiment, bearing structures 38 include
at least a #1 bearing structure 38-1 forward of the geared
architecture 48 and a #2 bearing structure 38-2 located aft of the
geared architecture 48.
[0033] A combustor 56 is arranged between the high pressure
compressor 52 and the high pressure turbine 54. A mid-turbine frame
57 of the engine static structure 36 is arranged generally between
the high pressure turbine 54 and the low pressure turbine 46. The
mid-turbine frame 57 can support one or more bearing structures 38
in the turbine section 28. The inner shaft 40 and the outer shaft
50 are concentric and rotate via the bearing structures 38 about
the engine centerline longitudinal axis A which is collinear with
their longitudinal axes. The inner shaft 40 and the outer shaft 50
can be either co-rotating or counter-rotating with respect to one
another.
[0034] The core airflow C is compressed by the low pressure
compressor 44 and the high pressure compressor 52, is mixed with
fuel and burned in the combustor 56, and is then expanded over the
high pressure turbine 54 and the low pressure turbine 46. The
mid-turbine frame 57 includes airfoils 59 which are in the core
airflow path. The high pressure turbine 54 and the low pressure
turbine 46 rotationally drive the respective high speed spool 32
and the low speed spool 30 in response to the expansion.
[0035] In some non-limiting examples, the gas turbine engine 20 is
a high-bypass geared aircraft engine. In a further example, the gas
turbine engine 20 bypass ratio is greater than about six (6:1). The
geared architecture 48 of the example gas turbine engine 20
includes an epicyclic gear train, such as a planetary gear system
or other gear system. The example epicyclic gear train has a gear
reduction ratio of greater than about 2.3. The geared architecture
48 enables operation of the low speed spool 30 at higher speeds
which can increase the operational efficiency of the low pressure
compressor 44 and low pressure turbine 46 and render increased
pressure in a fewer number of stages.
[0036] The low pressure turbine 46 pressure ratio is pressure
measured prior to inlet of low pressure turbine 46 as related to
the pressure at the outlet of the low pressure turbine 46 prior to
an exhaust nozzle of the gas turbine engine 20. In one non-limiting
embodiment, the bypass ratio of the gas turbine engine 20 is
greater than about ten (10:1), the fan diameter is significantly
larger than that of the low pressure compressor 44, and the low
pressure turbine 46 has a pressure ratio that is greater than about
5 (5:1). The geared architecture 48 of this embodiment is an
epicyclic gear train with a gear reduction ratio of greater than
about 2.5:1. It should be understood, however, that the above
parameters are only exemplary of one embodiment of a geared
architecture engine and that the present disclosure is applicable
to other gas turbine engines including direct drive turbofans.
[0037] In this embodiment of the example gas turbine engine 20, a
significant amount of thrust is provided by a bypass flow B due to
the high bypass ratio. The fan section 22 of the gas turbine engine
20 is designed for a particular flight condition--typically cruise
at about 0.8 Mach and about 35,000 feet. This flight condition,
with the gas turbine engine 20 at its best fuel consumption, is
also known as bucket cruise Thrust Specific Fuel Consumption
(TSFC). TSFC is an industry standard parameter of fuel consumption
per unit of thrust.
[0038] Fan Pressure Ratio is the pressure ratio across a blade of
the fan section 22 without the use of a Fan Exit Guide Vane system.
The low Fan Pressure Ratio according to one non-limiting embodiment
of the example gas turbine engine 20 is less than 1.45.
[0039] Low Corrected Fan Tip Speed is the actual fan tip speed
divided by an industry standard temperature correction of
"T"/518.7.sup.0.5. T represents the ambient temperature in degrees
Rankine. The Low Corrected Fan Tip Speed according to one
non-limiting embodiment of the example gas turbine engine 20 is
less than about 1150 fps (351 m/s).
[0040] FIG. 2 illustrates a portion 100 of a gas turbine engine,
such as the gas turbine engine 20. The portion 100 can include one
or more bearing structures 38. Only one bearing structure 38 is
depicted in FIG. 2 to schematically illustrate its features, but
this is in no way intended to limit this disclosure.
[0041] The bearing structure 38 supports a shaft 61, such as the
inner shaft 40 or the outer shaft 50, which supports a rotor
assembly 63, such as a rotor assembly of the compressor section 24
or the turbine section 28, through a hub 65. The rotor assembly 63
carries at least one airfoil 67 for adding or extracting energy
from the core airflow.
[0042] The bearing structure 38 defines a bearing compartment BC
that houses one or more bearings 71. The bearing compartment BC
contains a lubricant for lubricating (and acting as a cooling
medium to) the bearings 71. One or more seals 73 (two shown)
contain the lubricant within the bearing compartment BC. The seals
73 of the bearing compartment BC must be pressurized to prevent the
lubricant from leaking out during certain ground and flight
conditions (both steady state and transient). A buffer system can
be used to communicate buffer supply air to the bearing compartment
BC in order to provide adequate pressurization of the seals 73
without exceeding material and/or lubricant temperature
limitations. Example buffer systems that can be used for this and
other purposes are detailed below.
[0043] FIG. 3 illustrates an example buffer system 60 that can
communicate buffer supply air 62 to a portion of the gas turbine
engine 20, such as to one or more bearing compartments BC. In this
example, bearing compartments BC-1, BC-2, BC-3, BC-4(a), BC-4(b)
and BC-5 can be fed with buffer supply air 62. The buffer supply
air 62 pressurizes the bearing compartments BC and can maintain the
bearing compartments BC at an acceptable temperature. Although the
example embodiment illustrates communication of the buffer supply
air 62 to multiple bearing compartments BC-1 through BC-5 to
provide adequate bearing compartment seal pressurization to prevent
lubricant leakage, buffer supply air 62 could be communicated to
only a single bearing compartment or could be communicated for
anti-icing, ventilation, cooling and other purposes.
[0044] The buffer system 60 includes a first bleed air supply 64
and a second bleed air supply 66. In other words, the buffer system
60 is a dual supply system. In the exemplary embodiment, the first
bleed air supply 64 is a low pressure bleed air supply and the
second bleed air supply 66 is a high pressure bleed air supply that
includes a pressure that is greater than the pressure of the first
bleed air supply 64.
[0045] The first bleed air supply 64 can be sourced from the fan
section 22, the low pressure compressor 44 or the high pressure
compressor 52. In the illustrated non-limiting example, the first
bleed air supply 64 is sourced from an upstream stage of the high
pressure compressor 52. However, the first bleed air supply 64
could be sourced from any location that is upstream from the second
bleed air supply 66. The second bleed air supply 66 can be sourced
from the high pressure compressor 52, such as from a middle or
downstream stage of the high pressure compressor 52. The second
bleed air supply 66 could also be sourced from the low pressure
compressor 44 or the fan section 22 depending on from where the
first bleed air supply 64 is sourced.
[0046] The buffer system 60 can also include a valve 68 that is in
communication with both the first bleed air supply 64 and the
second bleed air supply 66. Although shown schematically, the first
bleed air supply 64 and the second bleed air supply 66 can be in
fluid communication with the valve 68 via buffer tubing, conduits,
or other passageways. Check valves may also be used to prevent the
second bleed air supply 66 from backflowing into the first bleed
air supply 64.
[0047] The valve 68 can select between the first bleed air supply
64 and the second bleed air supply 66 to communicate the buffer
supply air 62 to a desired portion(s) of the gas turbine engine 20.
In other words, the buffer supply air 62 that is communicated is
either the first bleed air supply 64 or the second bleed air supply
66 depending on which air supply is ultimately selected by the
valve 68, as is further discussed below.
[0048] The determination of whether to communicate the first bleed
air supply 64 or the second bleed air supply 66 as the buffer
supply air 62 is based on a power condition of the gas turbine
engine 20. The term "power condition" as used in this disclosure
generally refers to an operability condition of the gas turbine
engine 20. Gas turbine engine power conditions can include low
power conditions and high power conditions. Example low power
conditions include, but are not limited to, ground operation,
ground idle and descent idle. Example high power conditions
include, but are not limited to, takeoff, climb, and cruise
conditions. It should be understood that other power conditions are
also contemplated as within the scope of this disclosure.
[0049] In one exemplary embodiment, the valve 68 communicates the
first bleed air supply 64 (which is a relatively lower pressure
bleed air supply) as the buffer supply air 62 in response to
identifying a high power condition of a gas turbine engine 20. The
second bleed air supply 66 (which is a relatively higher pressure
bleed air supply) is selected by the valve 68 and communicated as
the buffer supply air 62 in response to detecting a low power
condition of the gas turbine engine 20. Both the first bleed air
supply 64 and the second bleed air supply 66 are intended to
maintain the same minimum pressure delta across the bearing
compartment seals. Low power conditions require a higher stage
pressure source to contain the lubricant within the bearing
compartment, while high power conditions require a lower stage
pressure source. The buffer system 60 can use the lowest possible
compressor stage to meet pressure requirements in order to minimize
supply temperature and any performance impact to the gas turbine
engine 20.
[0050] The valve 68 can be a passive valve. A passive valve
operates like a pressure regulator that can switch between two or
more sources without being commanded to do so by a controller, such
as an engine control (EEC). The valve 68 of this example uses only
a single input which is directly measured to switch between the
first bleed air supply 64 and the second bleed air supply 66.
[0051] The valve 68 could also be a controller based valve. For
example, the buffer system 60 can include a controller 70 in
communication with the valve 68 for selecting between the first
bleed air supply 64 and the second bleed air supply 66. The
controller 70 is programmed with the necessary logic for selecting
between the first bleed air supply 64 and the second bleed air
supply 66 in response to detecting a pre-defined power condition of
the gas turbine engine 20. The controller 70 could also be
programmed with multiple inputs.
[0052] In one example, a sensor 99 detects a power condition of the
gas turbine engine 20 and communicates a signal to the controller
70 to command modulation of the valve 68 between the first bleed
air supply 64 and the second bleed air supply 66. The valve 68
could also be modulated to an intermediate level to intermix the
first bleed air supply 64 and the second bleed air supply 66. Of
course, this view is highly schematic. It should be understood that
the sensor 99 and the controller 70 can be programmed to detect any
power condition. Also, the sensor 99 can be replaced by any control
associated with the gas turbine engine 20 or an associated
aircraft. Also, although shown as a separate feature, the
controller functionality could be incorporated into the valve
68.
[0053] FIG. 4 illustrates another example buffer system 160 that
can communicate buffer supply air 162 to provide adequate bearing
compartment seal pressurization at an acceptable temperature. The
buffer supply air 162 can also be used for additional purposes such
as anti-icing and ventilation or for other cooling requirements of
the gas turbine engine 20.
[0054] The buffer system 160 includes a first bleed air supply 164,
a second bleed air supply 166 and an ejector 172. If necessary, the
first bleed air supply 164 can be augmented by the ejector 172 to
prepare the buffer supply air 162 for communication to a portion of
the gas turbine engine 20, such as a bearing compartment BC
(schematically shown by FIG. 4). In other words, the ejector 172
can add pressure (using a relatively small amount of the second
bleed air supply 166) to the first bleed air supply 164 to prepare
the buffer supply air 162 for communication to an appropriate
location of a gas turbine engine 20. In one exemplary embodiment,
the ejector 172 can mix the first bleed air supply 164 of a first
pressure with the second bleed air supply 166 of a second higher
pressure to render the buffer supply air 162 of an intermediate
pressure to the first bleed air supply 164 and the second bleed air
supply 166.
[0055] The second bleed air supply 166, which is a higher pressure
air than the first bleed air supply 164, can be communicated to the
ejector 172 to power the ejector 172. The first bleed air supply
164 can be sourced from the fan section 22, the low pressure
compressor 44 or the high pressure compressor 52. The second bleed
air supply 166 can be sourced from a middle or downstream stage of
the high pressure compressor 52, or can include diffuser air. The
second bleed air supply 166 could also be sourced from the low
pressure compressor 44 or the fan section 22 depending on from
where the first bleed air supply 164 is sourced.
[0056] Augmentation of the first bleed air supply 164 prepares the
buffer supply air 162 at an adequate pressure and temperature to
pressurize the bearing compartment(s) BC. The determination of
whether or not to augment the first bleed air supply 164 with the
ejector 172 is based on a power condition of the gas turbine engine
20. Gas turbine engine power conditions can include low power
conditions and high power conditions. Example low power conditions
include, but are not limited to, ground operation, ground idle and
descent idle. Example high power conditions include, but are not
limited to, takeoff, climb, and cruise conditions. It should be
understood that other power conditions are also contemplated as
within the scope of this disclosure.
[0057] In one example, the first bleed air supply 164 is augmented
by the ejector 172 in response to detecting a low power condition
of the gas turbine engine 20 in order to communicate a buffer
supply air 162 having adequate pressurization. The amount of
augmentation performed on the first bleed air supply 164 can vary
depending upon the type of power condition that is detected and the
pressure requirements of the bearing compartment(s) BC. For
example, in one embodiment, the first bleed air supply 164 is not
augmented by the ejector 172 in response to detection of a high
power condition of the gas turbine engine 20. In other words, the
first bleed air supply 164 can be communicated as the buffer supply
air 162 without any augmentation in response to some power
conditions.
[0058] The buffer system 160 can include a controller 170 in
communication with the ejector 172 for determining whether or not
to augment the first bleed air supply 164. The controller 170 is
programmed with the necessary logic for making this determination
in response to detecting a pre-defined power condition of the gas
turbine engine 20. In one example, a sensor 199 detects a power
condition of the gas turbine engine 20 and communicates a signal to
the controller 170 to command the ejector 172 to augment the first
bleed air supply 64. Of course, this view is highly schematic. It
should be understood that the sensor 199 and the controller 170 can
be programmed to detect any power condition. Also, the sensor 199
can be replaced by any control associated with the gas turbine
engine 20 or an associated aircraft. Also, although shown as a
separate feature, the controller 170 functionality could be
incorporated into the ejector 172.
[0059] FIG. 5 illustrates yet another example buffer system 260. In
this example, the buffer system 260 is a two-circuit, multi-source
buffer system that includes at least a first circuit 274 and a
second circuit 276. Additional circuits could also be incorporated.
Low pressure requirements of the gas turbine engine 20 can be fed
with a first buffer supply air 262A from the first circuit 274,
while high pressure requirements of the gas turbine engine 20 can
be buffered with a second buffer supply air 262B from the second
circuit 276. In other words, the first circuit 274 can buffer a
first portion(s) of the gas turbine engine 20, while the second
circuit 276 can buffer a second, different portion(s). Example
components subject to low pressure requirements include bearing
compartments in low pressure regions of the gas turbine engine 20,
such as front or rear bearing compartments. Example components
subject to high pressure requirements include bearing compartments
in high pressure regions of the gas turbine engine 20, such as
mid-engine bearing compartments.
[0060] In this example, the first circuit 274 is similar to the
buffer system 60 of FIG. 3 and includes a first bleed air supply
264A, a second bleed air supply 266A and a valve 268A. The second
circuit 276 includes a first bleed air supply 264B, a second bleed
air supply 266B, a valve 268B and a conditioning device 280. In
this non-limiting example, the conditioning device 280 cools the
second buffer supply air 262B to an acceptable temperature for
addressing higher pressure requirements. The conditioning device
could include an air-to-air heat exchanger, a fuel-to-air heat
exchanger, or any other suitable heater exchanger. The conditioning
device 280 could also be a device other than a heat exchanger.
[0061] The second bleed air supply 266A of the first circuit 274
can be common to the first bleed air supply 264B of the second
circuit 276. These sources can also be completely separate. In each
of the first circuit 274 and the second circuit 276, the second
bleed air supplies 266A, 266B are communicated as the buffer supply
airs 262A, 262B for low power conditions of the gas turbine engine
20 and the first bleed air supplies 264A, 264B are communicated as
the buffer supply airs 262A, 262B in response to high power
conditions of the gas turbine engine 20. Example low power
conditions include, but are not limited to, ground operation,
ground idle, and flight idle conditions. Example high power
conditions include, but are not limited to, takeoff, climb, and
cruise conditions. It should be understood that other power
conditions are also contemplated as within the scope of this
disclosure.
[0062] In one exemplary embodiment, the valves 268A, 268B select
and communicate the first bleed air supplies 264A, 264B (which are
relatively lower pressure bleed air supplies) as the buffer supply
airs 262A, 262B in response to identifying a high power condition
of a gas turbine engine 20. The second bleed air supplies 266A,
266B (which are relatively higher pressure bleed air supplies) are
selected by the valves 268A, 268B and communicated as the buffer
supply airs 262A, 262B in response to detecting a low power
condition of the gas turbine engine 20. Both the lower bleed air
supplies and the higher bleed air supplies are intended to maintain
the same minimum pressure delta across the bearing compartment
seals. Low power conditions require a higher stage pressurize
source to contain the lubricant within the bearing compartment,
while high power conditions require a lower pressure stage source.
The buffer system 260 can use the lowest possible compressor stage
to meet the pressure requirements in order to minimize supply
temperature and any performance impact to the gas turbine engine
20.
[0063] The buffer system 260 can also include a controller 270 in
communication with the valves 268A, 268B for selectively switching
between the first bleed air supplies 264A, 264B and the second
bleed air supplies 266A, 266B. A single controller or multiple
controllers could be utilized. The controller 270 can also command
operation of the conditioning device 280 of the second circuit 276
for cooling the buffer supply air 262B. Alternatively, separate
controllers can be used to control each of the first circuit 274,
the second circuit 276 and the conditioning device 280.
[0064] Although the different examples have a specific component
shown in the illustrations, embodiments of this disclosure are not
limited to those particular combinations. It is possible to use
some of the components or features from one of the examples in
combination with features or components from another one of the
examples.
[0065] Furthermore, the foregoing description shall be interpreted
as illustrative and not in any limiting sense. A worker of ordinary
skill in the art would understand that certain modifications could
come within the scope of this disclosure. For these reasons, the
following claims should be studied to determine the true scope and
content of this disclosure.
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